EP2963250B1 - Revêtement pour isoler des composants métalliques de composants composites - Google Patents

Revêtement pour isoler des composants métalliques de composants composites Download PDF

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Publication number
EP2963250B1
EP2963250B1 EP15172924.1A EP15172924A EP2963250B1 EP 2963250 B1 EP2963250 B1 EP 2963250B1 EP 15172924 A EP15172924 A EP 15172924A EP 2963250 B1 EP2963250 B1 EP 2963250B1
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EP
European Patent Office
Prior art keywords
barrier coating
oxide
layer
silicon
coating
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EP15172924.1A
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German (de)
English (en)
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EP2963250A1 (fr
Inventor
Sean Landwehr
Sungbo Shim
Adam Chamberlain
Ann Bolcavage
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Rolls Royce Corp
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Rolls Royce Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/28—Supporting or mounting arrangements, e.g. for turbine casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/40—Heat treatment
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/90—Mounting on supporting structures or systems
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
    • F05D2300/131—Molybdenum
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/21—Oxide ceramics
    • F05D2300/2112—Aluminium oxides
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/22—Non-oxide ceramics
    • F05D2300/222—Silicon
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/611—Coating
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/70—Treatment or modification of materials
    • F05D2300/701—Heat treatment

Definitions

  • the present disclosure relates generally to gas turbine engines, and more specifically to coatings used in gas turbine engine assemblies.
  • Gas turbine engine components are exposed to high temperature environments with an increasing demand for even higher temperatures. Economic and environmental concerns relating to the reduction of emissions and the increase of efficiency are driving the demand for higher gas turbine operating temperatures. In order to meet these demands, temperature capability of the components in hot sections such as blades, vanes, blade tracks, and combustor liners must be increased.
  • Ceramic matrix composites may be a candidate for inclusion in the hot sections where higher gas turbine engine operating temperatures are required.
  • One benefit of ceramic matrix composite engine components is the high-temperature mechanical, physical, and chemical properties of the ceramic matrix composite components which allow the gas turbine engines to operate at higher temperatures than current engines.
  • the ceramic matrix composite components may be held in place by metallic structures.
  • the metallic structures may interact chemically with the ceramic matrix composite at high temperatures when used over long durations. In some cases, the interaction of metallic structures and ceramic matrix composites supported thereon, may lead to degradation of the metallic structures.
  • EP 1693478 relates to a method for inhibiting diffusion of silicon into a support structure from a component formed of a silicon-containing material.
  • a method of isolating a metallic support component from a silicon-comprising composite component in a gas turbine engine may include applying a precursor coating onto the metallic support component, mounting the silicon-comprising composite component so that the silicon-comprising composite engages the precursor coating applied to the metallic support component to form an engine assembly, and operating the gas turbine engine comprising the engine assembly so that the precursor coating is heated to a predetermined temperature to form a dual layer barrier coating comprising an oxide layer along an exterior edge of a base layer from the precursor coating so that the silicon in the silicon-comprising composite component is restricted from ingress into the metallic support component by the oxide-comprising layer during further operation of the gas turbine engine.
  • the precursor coating includes a refractory metal that assists the formation of the oxide-comprising layer.
  • the refractory metal included in the precursor coating is selected from the group consisting of molybdenum, tungsten, and tantalum.
  • the precursor coating comprising between about 1 weight percent and about 60 weight percent of the refractory metal.
  • the precursor coating may comprise an oxide selected from the group consisting of chromium oxide, aluminum oxide, and silicon oxide.
  • the oxide-comprising layer of the barrier coating may have a thickness of between about 0.5 microns and about 10 microns.
  • the barrier coating may have a thickness of between about 25 microns and about 300 microns.
  • the temperature that causes the formation of the oxide-comprising layer along an exterior edge of the base layer of the barrier coating is between about 800°C (1,500°F) and about 1000°C (1,800°F).
  • the precursor coating may comprise an oxide selected from the group consisting of chromium oxide, aluminum oxide, and silicon oxide.
  • the precursor coating includes a refractory metal that assists the formation of the oxide-comprising layer.
  • the refractory metal included in the precursor coating is selected from the group consisting of molybdenum, tungsten, and tantalum. In some embodiments the precursor coating may comprise between about 1 weight percent and about 60 weight percent of the refractory metal.
  • the oxide-comprising layer of the barrier coating may have a thickness of between about 0.5 microns and about 10 microns.
  • the barrier coating may have a thickness of between about 25 microns and about 300 microns.
  • the temperature that causes the formation of the oxide-comprising layer along an exterior edge of the base layer of the barrier coating may be between about 800°C (1,500°F) and about 1000°C (1,800°F).
  • an engine assembly for use in a gas turbine engine.
  • the engine assembly may include a metallic support component, a silicon-comprising composite component mounted to the metallic support component so that the hanger supports the ceramic matrix composite component, and a barrier coating on the metallic support component so that the silicon-comprising component engages the barrier coating without contacting the metallic support component, the barrier coating comprising an interior base layer and an exterior oxide-comprising layer that is engaged by the silicon-comprising composite component, the exterior oxide layer having a thickness of between about 0.5 microns and about 15 microns.
  • the barrier coating comprises between about 1 weight percent and about 60 weight percent of a refractory metal to assist in formation of an exterior oxide-comprising layer upon heating the barrier coating to a predetermined temperature when the engine assembly is used in a gas turbine engine.
  • the refractory metal included in the barrier coating is selected from the group consisting of molybdenum, tungsten, and tantalum.
  • the barrier coating comprising between about 1 weight percent and about 60 weight percent of the refractory metal.
  • the barrier coating may comprise an oxide selected from the group consisting of chromium oxide, aluminum oxide, and silicon oxide.
  • the metallic support component may comprise a hanger which may include a radially-extending portion and an axially-extending portion that extends from the radially-extending portion.
  • the barrier coating may be applied to the axially-extending portion, and the barrier coating may have a thickness that decreases as the axially-extending portion extends away from the radially-extending portion.
  • the exterior oxide-comprising layer may be formed by a process comprising the steps of (i) assembling the metallic hanger and the silicon-comprising composite component into a gas turbine engine and (ii) heating a precursor coating applied to the metallic hanger at the interface of the metallic hanger with the silicon-comprising component to a predetermined temperature.
  • An illustrative aerospace gas turbine engine 10 may include an output shaft 12, a compressor section 14, a combustor section 16, and a turbine section 18 all mounted to a case 20 as shown in Fig. 1 .
  • the output shaft 12 may be coupled to a propeller (not shown) and may be driven by the turbine section 18.
  • the compressor section 14 may compress and deliver air to the combustor section 16.
  • the combustor section16 may mix fuel with the compressed air received from the compressor section 14 to ignite the fuel.
  • the hot high pressure products of the combustion reaction in the combustor section 16 may be directed into the turbine section 18 and the turbine section 18 may extract work to drive the compressor section 14 and the output shaft 12, as suggested in Fig. 1 .
  • the turbine section 18 illustratively may include static turbine vane assemblies 21, 22 and corresponding turbine wheel assemblies 24, 25 as shown in Fig. 1 .
  • Each vane assembly 21, 22 may include a plurality of corresponding vanes 26, 27, etc. and each turbine wheel assembly 24, 25 may include a plurality of corresponding blades 28, 29.
  • the vanes 26, 27 of the vane assemblies 21, 22 may extend across the flow path of the hot, high-pressure combustion products from the combustor 16 to direct the combustion products toward the blades 28, 29 of the turbine wheel assemblies 24, 25.
  • the blades 28, 29 may in turn be pushed by the combustion products to cause the turbine wheel assemblies 26, 27 to rotate; thereby, driving the rotating components of the compressor section 14 and the output shaft 12.
  • the turbine section 18 also includes a plurality of turbine shrouds 30, 31 that extend around each turbine wheel assembly 24, 25 to block combustion products from passing over the blades 28, 29 without pushing the blades 28, 29 to rotate, as suggested in Fig 1 .
  • the turbine shroud 30 may include a carrier 32 and a blade track (sometimes called a seal ring) 34, as shown in Figs. 2 and 3 .
  • the carrier 32 may be an annular, round, metallic component and may support the blade track 34 in position adjacent to the blades 28 of the turbine wheel assembly 24.
  • the illustrative blade track 34 may be made from silicon-comprising ceramic matrix composite materials.
  • the blade track 34 may include a retainer 52 that engages the carrier 32 to position the blade track 34 relative to other static turbine components in the gas turbine engine 10.
  • a barrier coating 60 may be adhered to the carrier 32 at interfaces of the carrier 32 with the blade track 34 as shown in Figs. 3 and 4 . As suggested, the barrier coating 60 may block the diffusion or ingress of silicon or other similar elements from the carrier 32 into the blade track 34.
  • the barrier coating 60 may include an oxide layer 62 along the exterior edge or exterior surface of a base layer 64.
  • the precursor coating 63 may be heated to a predetermined temperature to cause formation of the oxide layer 62 along an exterior surface of the base layer 64 as suggested in Figs. 5 and illustratively shown in Figs 6-10 .
  • the base layer 64 may be sandwiched between the oxide layer 62 and the metallic hanger 44 to create the dual-layer coating.
  • the base layer 64 substantially may not include an oxidant and may not be further oxidized.
  • the base layer is essentially the same as precursor coating.
  • An area between the oxide layer 62 and the base layer 64 may not be substantially distinct and may include overlapping of the oxide layer 62 and the base layer 64. Once the oxide layer 62 is formed, the interface between the oxide layer 62 and the base layer 64 may be distinct.
  • the precursor coating 63 may be heated during use of the turbine shroud 30 in the gas turbine engine 10 to cause formation of the oxide layer 62 exterior to the base layer 64, as shown in Figs. 6-10 .
  • heat treatment of the turbine shroud 30 in a furnace may be performed prior to use in the gas turbine engine 10 to cause formation of the oxide layer 62 exterior to the base layer 64.
  • the barrier coating 60 and/or the oxide layer 62 may include chromium oxide, aluminum oxide, and/or silicon oxide.
  • a refractory metal such as molybdenum, tungsten, and/or tantalum is included in the barrier coating 60 to assist in the formation of the oxide layer 62 during a heating process.
  • the illustrative barrier coating 60 may have a thickness T of between about 25 microns and about 300 microns, as depicted in Fig.4 .
  • the oxide layer 62 of the barrier coating 60 may have a thickness t of between about 0.6 microns and about 10 microns, as depicted in Fig. 4 .
  • the barrier coating 60 may have an axial thickness that decreases as the axially-extending portions 43, 45 extend away from the radially-extending portions 41, 46 of the forward and aft hangers 42, 44 as depicted in Fig 4 .
  • the barrier coating 60 is thicker at the locations adjacent to the radially-extending portions 41, 46 than at locations spaced apart from the radially-extending portions 41, 46. This arrangement may reduce forces applied to the carrier 32 from the blade track 34.
  • the carrier 32 may include an attachment flange 38 coupled to the case 20, a forward hanger 42, and an aft hanger 44, as shown in Figs. 2-4 .
  • the forward hanger 42 illustratively may have a radially-extending portion 41 and an axially-extending portion 43 for hanging the blade track 34.
  • the aft hanger 44 like the forward hanger 42, illustratively may have a radially-extending portion 46 and an axially-extending portion 45 for hanging the blade track 34, as shown in Figs. 3-4 .
  • the carrier 32 may be made from a metallic alloy such as nickel-based or cobalt-based alloy.
  • the blade track 34 may include a runner 48, a forward attachment arm 50 and an aft attachment arm 54, as shown in Figs 2-4 .
  • the runner 48 may extend around the turbine wheel assembly 24 to block gasses from passing over the turbine blades 28 without pushing the blades 28.
  • the forward attachment arm 50 may have a radially-extending portion 51 and may have an axially-extending portion 52.
  • the aft attachment arm 54 may have a radially-extending portion 55 and an axially-extending portion 56 for attaching to the carrier 32.
  • the blade track 34 may include or be formed of a silicon-carbide/silicon-carbide ceramic matrix composite.
  • the silicon-comprising blade track 34 may interact with the nickel or any number of constituent materials of the metallic carrier 32 Free Si from the composite diffuses into the metallic component and may react with nickel and other alloy elements, which may degrade performance of the component. Silicon in large quantities may alloy with the metallic carrier and may form phases having a lower melting point than the surrounding nickel-based material. The interaction may degrade the properties and performance of the carrier 32, if direct contact between the components is allowed.
  • the barrier coating 60 may reduce silicon diffusion and other reactions allowing the carrier 32 to retain desired properties.
  • the barrier coating 60 may be applied to the axially-extending portions 43, 45 of the forward and aft hangers 42, 44 as shown in Fig. 3 .
  • the barrier coating 60 on the axially-extending portions 43, 45, of the forward and aft hangers 42, 44 may mate with or engage with the blade track 34. While in the preceding example the barrier coating 60 is shown and described in conjunction with the turbine shroud 30, it may be incorporated at other interfaces throughout the gas turbine engine 10. More specifically, the barrier coating 60 may be used at the interface of any metallic component with a composite component to block chemical interaction between the metallic component and the composite component.
  • metallic combustor supports may hold composite liner tiles in place and the barrier coating 60 may be applied at the interface between the metallic combustor supports and the composite liner tiles.
  • metallic turbine rotors may hold composite turbine blades in place around turbine wheels and the barrier coating 60 may be applied at the interface between the metallic turbine rotors and the composite turbine blades.
  • a metallic support component such as the carrier 32
  • a composite component such as a blade track 34
  • a metallic support component is provided for mounting a composite component in a gas turbine engine 10 as suggested in Figs. 5 shown illustratively with the metallic aft hanger 44 in Fig. 6 .
  • a precursor coating 63 may be applied to the ceramic mating surface of a metallic component such as the aft hanger 44 of the carrier 32 as suggested in Fig. 5 and illustratively depicted in Fig. 7 .
  • the precursor coating 63 may be transformed into the dual-layer barrier coating 60 which may include a base layer 64 and an oxide layer 62.
  • the precursor coating 63 may be applied as a single layer or a plurality of layers using electro-deposition, chemical vapor deposition, physical vapor deposition or any other suitable process as depicted in Fig. 7 .
  • the precursor coating 63 may include chromium, aluminum, silicon, cobalt, nickel and/or any other alloys.
  • the precursor coating may be Co-based and may need Cr and/or Al to form chromium or aluminium on the coating surface. Nickel may be added for improved oxidation resistance, but silicon may easily diffuse into nickel.
  • the precursor coating 63 may include a refractory metal such as molybdenum, tungsten, and/or tantalum, which may act as a silicon getter.
  • the composite component may be mounted to the metallic support component and assembled in a gas turbine engine so the precursor coating 63 is engaged, as suggested in Fig. 5 and illustratively depicted in Fig. 8 .
  • the gas turbine engine 10 may be operated to heat the precursor coating 63 forming an oxide layer 62 exterior to a base layer 64 within the barrier coating 60, as suggested in Fig. 5 and illustratively depicted in Fig. 9 . Operating the gas turbine engine 10 to temperatures between about 1,500 degrees Fahrenheit and about 1,800 degrees Fahrenheit in an atmosphere that includes oxygen will cause the formation of the oxide layer on the exterior surface of the base layer.
  • the precursor coating 63 may be heat treated to a predetermined temperature that may cause the formation of an oxide layer 62 exterior to the base layer 64 prior to mounting the composite component on the metallic support component in the gas turbine engine 10.
  • the heating of the precursor coating 63 may allow for the formation of the oxide layer 62 on the exterior edge of the barrier coating 60 creating the dual layer coating wherein the base layer is sandwiched between the oxide layer 62 and the metallic aft hanger 44.
  • the barrier coating 60 may be reheated through engine operation to seal and repair the coating.
  • the damaged portions of the oxide layer 62 may be sealed after normal use of the assembly, as depicted in Fig. 10 .
  • the oxide layer 62 separates from the base layer 64
  • the base layer 64 that may be exposed from the loss of the oxide layer 62 may oxidize to form a further oxidized layer, as depicted in Fig. 10 .
  • another layer or a plurality of layers of precursor coating 63 may be added and heated to create the dual layer barrier coating 60 to repair any cracks in the barrier coating.
  • Heating the precursor coating 63 may occur during engine 10 operation and/or during a heat treatment applied before assembly in a gas turbine engine 10.
  • the precursor coating 63 may include chromium, aluminum, silicate and/or other materials.
  • the precursor coating 63 may also include a refractory metal that makes up between about 0.1 and about 60 weight percent and may assist in the formation of the exterior oxide comprising layer.
  • the following examples are illustrative of the invention and are not intended to limit the scope of the invention.
  • the precursor coating 63 may be applied to the forward hanger 42 and aft hanger 44 of the carrier 32 using thermal spray coating and/or air plasma spray.
  • the precursor coating 63 may be applied to the forward hanger 42 and aft hanger 44 of the carrier 32 using thermal spray coating and/or air plasma spray. Heat treatment of the Amdry 995C precursor coating 63 for 100 hours at 1600 degrees Fahrenheit, while in contact with a silicon-comprising composite component, resulted in no diffusion of silicon into the base layer 364 of the barrier coating 360 as shown in Fig 7 .
  • a precursor coating 63 of Amdry MM509 received from Sulzer-Metco (Co 23.Cr 10 Ni 7W 3.5Ta 06.C weight %) was heated to form a barrier coating 460 comprising a base layer 464 and an exterior oxide layer 462 as shown in the cross sectioning of the barrier coating 460 in Fig. 13 .
  • the precursor coating 63 may be applied to the forward hanger 42 and aft hanger 44 of the carrier 32 using thermal spray coating and/or air plasma spray.
  • Heat treatment of the Amdry MM509 precursor coating 63 for 100 hours at 1600 degrees Fahrenheit while in contact with a silicon-comprising composite component resulted in no diffusion of silicon into the base layer 464 of the barrier coating 460 as shown in Fig 8 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Other Surface Treatments For Metallic Materials (AREA)

Claims (10)

  1. Méthode d'isolation d'un élément de support métallique (32) d'un élément composite comprenant du silicium (34) dans un moteur à turbine à gaz (10), la méthode comprenant
    l'application d'un revêtement précurseur (63) sur l'élément de support métallique,
    le montage de l'élément composite comprenant du silicium de sorte que l'élément composite comprenant du silicium engage le revêtement précurseur appliqué sur l'élément de support métallique pour former un ensemble moteur, et
    le chauffage du revêtement précurseur à une température prédéterminée pour constituer un revêtement protecteur à double couche sur l'élément de support métallique, le revêtement protecteur à double couche comprenant une couche d'oxyde (62) le long d'un bord extérieur d'une couche de base (64) depuis le revêtement précurseur, de façon à restreindre l'introduction de silicium dans l'élément composite comprenant du silicium dans l'élément de support métallique par la couche comprenant de l'oxyde au cours de l'utilisation ultérieure du moteur à turbine à gaz,
    caractérisée en ce que :
    le revêtement précurseur comprend entre 1 pour cent en poids et 60 pour cent en poids d'un métal réfractaire sélectionné dans le groupe composé de molybdène, de tungstène, et de tantale.
  2. Méthode selon la revendication 1, la couche comprenant de l'oxyde comprenant un oxyde sélectionné dans le groupe composé d'oxyde de chrome, d'oxyde d'aluminium et d'oxyde de silicium.
  3. Méthode selon la revendication 2, la couche de précurseur comprenant un métal de base sélectionné dans le groupe composé de nickel, de cobalt, et d'aluminium.
  4. Méthode selon une quelconque des revendications 1-3, l'épaisseur de la couche comprenant de l'oxyde du revêtement protecteur étant comprise entre environ 0,5 micron et environ 15 microns.
  5. Méthode selon la revendication 4, l'épaisseur du revêtement protecteur étant comprise entre environ 25 microns et environ 300 microns.
  6. Méthode selon une quelconque des revendications 1-5, la température prédéterminée causant la formation de la couche comprenant de l'oxyde le long d'un bord extérieur de la couche de base du revêtement protecteur étant comprise entre environ 800°C et environ 1 000°C.
  7. Méthode selon une quelconque des revendications 1 à 6, la température prédéterminée étant atteinte au cours de l'utilisation du moteur à turbine à gaz.
  8. Ensemble moteur pour un moteur à turbine à gaz, l'ensemble comprenant
    un élément de support métallique (32),
    un élément composite comprenant du silicium (34) monté sur l'élément de support métallique de sorte que l'élément de support métallique supporte l'élément composite comprenant du silicium, et
    un revêtement protecteur (60) sur l'élément de support métallique de sorte que l'élément composite comprenant du silicium engage le revêtement protecteur sans contacter l'élément de support métallique, le revêtement protecteur comprenant une couche de base intérieure (64) et une couche extérieure comprenant de l'oxyde (62) engagée par l'élément composite comprenant du silicium,
    l'épaisseur de la couche extérieure comprenant de l'oxyde étant comprise entre environ 0,5 micron et environ 15 microns,
    caractérisé en ce que :
    le revêtement protecteur comprend un métal réfractaire dans la mesure de 1 pour cent en poids à 60 pour cent en poids choisi parmi le groupe constitué de molybdène, de tungstène, et de tantale pour contribuer à la formation d'une couche extérieure comprenant de l'oxyde lors du chauffage du revêtement protecteur à une température prédéterminée lors de l'utilisation de l'ensemble moteur dans un moteur à turbine à gaz.
  9. Ensemble moteur selon la revendication 8, l'élément de support métallique comprenant une partie à extension radiale et une partie à extension axiale déployées depuis la partie s'étendant radialement, le revêtement protecteur étant appliqué sur la partie à extension axiale, et l'épaisseur axiale du revêtement protecteur diminuant au fur et à mesure que la partie à extension axiale s'éloigne de la partie à extension radiale.
  10. Ensemble moteur selon la revendication 8 ou la revendication 9, l'élément de support métallique comprenant un dispositif de suspension métallique.
EP15172924.1A 2014-06-30 2015-06-19 Revêtement pour isoler des composants métalliques de composants composites Active EP2963250B1 (fr)

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CA2895629A1 (fr) 2015-12-30
US20150377069A1 (en) 2015-12-31
US9920656B2 (en) 2018-03-20

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